Turbine engine with airfoil having high acceleration and low blade turning

ABSTRACT

A turbine engine with at least a compressor section, combustor section, turbine section and a set of airfoils. The airfoils include geometric characteristics to create a high contraction ratio (CR), a low blade turning (BT) at a radially inward location the airfoil, a low solidity, or a low aspect ratio (AR).

CROSS REFERENCE TO RELATED APPLICATIONS

This application is a continuation of U.S. patent application Ser. No.17/148,635, filed Jan. 14, 2021, now allowed, which claims priority toItalian Patent Application No. 102020000005146, filed Mar. 11, 2020,both of which is incorporated herein by reference in its entirety.

TECHNICAL FIELD

The disclosure generally relates to an airfoil for an engine, and morespecifically to the geometry of said airfoil.

BACKGROUND

Turbine engines, and particularly gas or combustion turbine engines, arerotary engines that extract energy from a flow of combusted gasespassing through the engine onto a multitude of rotating turbine blades.

A turbine engine includes but is not limited to, in serial flowarrangement, a forward fan assembly, an aft fan assembly, ahigh-pressure compressor for compressing air flowing through the engine,a combustor for mixing fuel with the compressed air such that themixture can be ignited, and a high-pressure turbine. The high-pressurecompressor, combustor and high-pressure turbine are sometimescollectively referred to as the core engine.

In at least some turbine engines, at least one turbine rotates in anopposite direction than the other rotating components within the engine.In some implementations a counter-rotating low-pressure turbine includesan outer drum having a first set of stages that are rotatably coupled tothe forward fan assembly, and an inner drum having an equal number ofstages that is rotatably coupled to the aft fan assembly.

Turbine engines include several components that utilize airfoils. By waya of non-limiting example, the airfoils can be located in the engineturbines, compressors, or fans. The geometry of the airfoil can affectvarious characteristics such as, but not limited to, contraction ratio,blade turning, solidity, or aspect ratio

BRIEF DESCRIPTION

In one aspect, the disclosure relates to a turbine engine comprising atleast one blade carried by a rotor and rotating about a rotational axis,the blade comprising, an outer wall defining a pressure side and asuction side extending in a chord-wise direction between a leading edgeto a trailing edge and extending in a span-wise direction between a rootand a tip, a mean camber line extending between a leading edge and atrailing edge and intersecting the leading edge to define a leading edgeintersection, and intersecting the trailing edge to define a trailingedge intersection, an inlet angle, β_(in), defined by an included anglebetween a line parallel to the mean camber line at the leading edgeintersection and the rotational axis, an outlet angle, β_(out), definedby an included angle between a line parallel to the mean camber line atthe trailing edge intersection and the rotational axis, wherein theblade has a contraction ratio (CR) of greater than 0.45 along at least80% of a span of the at least one blade, where the CR is determined bythe formula:

${CR} = {1 - ( \frac{\cos( \beta_{out} )}{\cos( \beta_{in} )} )}$

wherein the blade has a blade turning (BT) of less than 100 degreesalong at least 30% of the span, where the blade turning is determined bythe formula:

${BT} = {{❘\beta_{out}❘} + \frac{\beta_{in}*\beta_{out}}{❘\beta_{out}❘}}$

In another aspect, the disclosure relates to an airfoil configured torotate about a rotational axis and comprising an outer wall defining apressure side and a suction side extending in a chord-wise directionbetween a leading edge to a trailing edge and extending in a span-wisedirection between a root and a tip, a mean camber line extending betweena leading edge and a trailing edge and intersecting the leading edge todefine a leading edge intersection, and intersecting the trailing edgeto define a trailing edge intersection, an inlet angle, β_(in), definedby an included angle between a line parallel to the mean camber line atthe leading edge intersection and the rotational axis, an outlet angle,β_(out), defined by an included angle between a line parallel to themean camber line at the trailing edge intersection and the rotationalaxis, wherein the airfoil has a contraction ratio (CR) of greater than0.45 along at least 80% of a span of the at least one blade, where theCR is determined by the formula:

${CR} = {1 - ( \frac{\cos( \beta_{out} )}{\cos( \beta_{in} )} )}$

wherein the airfoil has a blade turning (BT) of less than 100 degreesalong at least 30% of the span, where the blade turning is determined bythe formula:

${BT} = {{❘\beta_{out}❘} + \frac{\beta_{in}*\beta_{out}}{❘\beta_{out}❘}}$

BRIEF DESCRIPTION OF THE DRAWINGS

In the drawings:

FIG. 1 is a schematic cross-sectional diagram for a gas turbine engine.

FIG. 2 is a perspective view of an airfoil that can be used with a gasturbine of FIG. 1 .

FIG. 3 is schematic view of a profile of the airfoil of FIG. 2 .

FIG. 4 is a schematic view of a plurality of airfoils of FIG. 2 placedon an engine component.

FIG. 5 is a side view of an airfoil of FIG. 2 placed on a rotatablecomponent.

DETAILED DESCRIPTION OF THE INVENTION

Aspects of this description are broadly directed to an airfoil with aunique profile with a predetermined contraction ratio (CR) andpredetermined blade turning (BT), which collectively provide the airfoilwith the ability to reduce boundary layer growth from the leading edgeto the trailing edge. This profile can be used in a wide range ofenvironments including environments having a high direct or relativerotational speed and higher centrifugal forces when compared toconventional turbine engines. Blade solidity or aspect ratio (AR) canalso be controlled to further enhance the ability of the unique profileto retard boundary layer growth. The unique profile can also reducemanufacturing and material costs associated with the turbine engine.

As used herein, the term “upstream” refers to a direction that isopposite the fluid flow direction, and the term “downstream” refers to adirection that is in the same direction as the fluid flow. The term“fore” or “forward” means in front of something and “aft” or “rearward”means behind something. For example, when used in terms of fluid flow,fore/forward can mean upstream and aft/rearward can mean downstream.While “a set of” various elements will be described, it will beunderstood that “a set” can include any number of the respectiveelements, including only one element.

Additionally, as used herein, the terms “radial” or “radially” refer toa direction away from a common center. For example, in the overallcontext of a turbine engine, radial refers to a direction along a rayextending between a center longitudinal axis of the engine and an outerengine circumference. Furthermore, as used herein, the term “set” or a“set” of elements can be any number of elements, including only one.

All directional references (e.g., radial, axial, proximal, distal,upper, lower, upward, downward, left, right, lateral, front, back, top,bottom, above, below, vertical, horizontal, clockwise, counterclockwise,upstream, downstream, forward, aft, etc.) are only used foridentification purposes to aid the reader's understanding of the presentdisclosure, and do not create limitations, particularly as to theposition, orientation, or use of aspects of the disclosure describedherein. Connection references (e.g., attached, coupled, secured,fastened, connected, and joined) are to be construed broadly and caninclude intermediate members between a collection of elements andrelative movement between elements unless otherwise indicated. As such,connection references do not necessarily infer that two elements aredirectly connected and in fixed relation to one another. The exemplarydrawings are for purposes of illustration only and the dimensions,positions, order and relative sizes reflected in the drawings attachedhereto can vary.

As used herein, the term “forward” or “upstream” refers to moving in adirection toward the engine inlet, or a component being relativelycloser to the engine inlet as compared to another component. The term“aft” or “downstream” used in conjunction with “forward” or “upstream”refers to a direction toward the rear or outlet of the engine or beingrelatively closer to the engine outlet as compared to another component.Additionally, as used herein, the terms “radial” or “radially” refer toa dimension extending between a center longitudinal axis of the engineand an outer engine circumference. Furthermore, as used herein, the term“set” or a “set” of elements can be any number of elements, includingonly one.

All directional references (e.g., radial, axial, proximal, distal,upper, lower, upward, downward, left, right, lateral, front, back, top,bottom, above, below, vertical, horizontal, clockwise, counterclockwise,upstream, downstream, forward, aft, etc.) are only used foridentification purposes to aid the reader's understanding of the presentdisclosure, and do not create limitations, particularly as to theposition, orientation, or use of aspects of the disclosure describedherein. Connection references (e.g., attached, coupled, connected, andjoined) are to be construed broadly and can include intermediate membersbetween a collection of elements and relative movement between elementsunless otherwise indicated. As such, connection references do notnecessarily infer that two elements are directly connected and in fixedrelation to one another. The exemplary drawings are for purposes ofillustration only and the dimensions, positions, order and relativesizes reflected in the drawings attached hereto can vary.

FIG. 1 is a schematic cross-sectional diagram of a turbine engine 10 foran aircraft. The turbine engine 10 has a generally longitudinal axis orcenterline 12 extending forward 14 to aft 16. The turbine engine 10includes, in downstream serial flow relationship, a fan section 18including a fan 20, a compressor section 22 including a booster or lowpressure (LP) compressor 24 and a high pressure (HP) compressor 26, acombustion section 28 including a combustor 30, a turbine section 32including a HP turbine 34, and a LP turbine 36, and an exhaust section38.

The fan section 18 includes a fan casing 40 surrounding the fan 20. Thefan 20 includes a plurality of fan blades 42 disposed radially about thecenterline 12. The HP compressor 26, the combustor 30, and the HPturbine 34 form an engine core 44 of the turbine engine 10, whichgenerates combustion gases. The engine core 44 is surrounded by corecasing 46, which can be coupled with the fan casing 40.

A HP shaft or spool 48 disposed coaxially about the centerline 12 of theturbine engine 10 drivingly connects the HP turbine 34 to the HPcompressor 26. A LP shaft or LP spool 50, which is disposed coaxiallyabout the centerline 12 of the turbine engine 10 within the largerdiameter annular HP spool 48, drivingly connects the LP turbine 36 tothe LP compressor 24 and fan 20. The HP and LP spools 48, 50 arerotatable about the engine centerline and couple to a plurality ofrotatable elements, which can collectively define an inner rotor/stator51. While illustrated as a root, it is contemplated that the innerrotor/stator 51 can be a stator.

The LP compressor 24 and the HP compressor 26 respectively include aplurality of compressor stages 52, 54, in which a set of compressorblades 56, 58 rotate relative to a corresponding set of staticcompressor vanes 60, 62 to compress or pressurize the stream of fluidpassing through the stage. In a single compressor stage 52, 54, multiplecompressor blades 56, 58 can be provided in a ring and can extendradially outwardly relative to the centerline 12, from a blade platformto a blade tip, while the corresponding static compressor vanes 60, 62are positioned upstream of and adjacent to the rotating compressorblades 56, 58. It is noted that the number of blades, vanes, andcompressor stages shown in FIG. 1 were selected for illustrativepurposes only, and that other numbers are possible.

The compressor blades 56, 58 for a stage of the compressor can bemounted to (or integral to) a disk 61, which is mounted to thecorresponding one of the HP and LP spools 48, 50. The static compressorvanes 60, 62 for a stage of the compressor can be mounted to the corecasing 46 in a circumferential arrangement.

The HP turbine 34 and the LP turbine 36 respectively include a pluralityof turbine stages 64, 66, in which a set of turbine blades 68, 70 arerotated relative to a corresponding set of static turbine vanes 72, 74(also called a nozzle) to extract energy from the stream of fluidpassing through the stage. In a single turbine stage 64, 66, multipleturbine blades 68, 70 can be provided in a ring and can extend radiallyoutwardly relative to the centerline 12 while the corresponding staticturbine vanes 72, 74 are positioned upstream of and adjacent to therotating blades 68, 70. It is noted that the number of blades, vanes,and turbine stages shown in FIG. 1 were selected for illustrativepurposes only, and that other numbers are possible.

The blades 68, 70 for a stage of the turbine can be mounted to a disk71, which is mounted to the corresponding one of the HP and LP spools48, 50. The static turbine vanes 72, 74 for a stage of the turbine canbe mounted to the core casing 46 in a circumferential arrangement.

Complementary to the rotor portion, the stationary portions of theturbine engine 10, such as the static compressor vanes and turbineblades 60, 62, 72, 74 among the compressor and turbine sections 22, 32are also referred to individually or collectively as an outerrotor/stator 63. As illustrated, the outer rotor/stator 63 can refer tothe combination of non-rotating elements throughout the turbine engine10. Alternatively, the outer rotor/stator 63 that circumscribes at leasta portion of the inner rotor/stator 51, can be designed to rotate.

In operation, the airflow exiting the fan section 18 is split such thata portion of the airflow is channeled into the LP compressor 24, whichthen supplies pressurized airflow 76 to the HP compressor 26, whichfurther pressurizes the air. The pressurized airflow 76 from the HPcompressor 26 is mixed with fuel in the combustor 30 and ignited,thereby generating combustion gases. Some work is extracted from thesegases by the HP turbine 34, which drives the HP compressor 26. Thecombustion gases are discharged into the LP turbine 36, which extractsadditional work to drive the LP compressor 24, and the exhaust gas isultimately discharged from the turbine engine 10 via the exhaust section38. The driving of the LP turbine 36 drives the LP spool 50 to rotatethe fan 20 and the LP compressor 24.

It is contemplated that a geared portion or a gear box can be includedwithin at least a portion of the turbine engine 10. The geared portioncan be configured to rotate one or more portions of the turbine engine10 at a desired rotational velocity. For example, the LP spool 50 can besegmented such that the portion of the LP spool 50 connected to the LPturbine 36 acts as an input to a gear box of the LP spool. The remainingportion of the LP spool 50 can act as an output from the gear box of theLP spool and be operatively coupled to the fan 20 and the LP compressor24. The gear box of the LP spool can be configured to provide a gearreduction between the LP turbine 36, the LP compressor 24, and the fan20. As such, the LP compressor 24 and the fan 20 can rotate a firstrotational velocity, while the LP turbine can rotate at a secondrotational velocity different than the first rotational velocity. Itwill be appreciated that this is non-limiting example and that thegeared portion can be applied to any suitable portion of the turbineengine 10.

A portion of the pressurized airflow 76 can be drawn from the compressorsection 22 as bleed air 77. The bleed air 77 can be drawn from thepressurized airflow 76 and provided to engine components requiringcooling. The temperature of pressurized airflow 76 entering thecombustor 30 is significantly increased. As such, cooling provided bythe bleed air 77 is necessary for operating of such engine components inthe heightened temperature environments.

A remaining portion of an airflow 78 bypasses the LP compressor 24 andengine core 44 and exits the turbine engine 10 through a stationary vanerow, and more particularly an outlet guide vane assembly 80, comprisinga plurality of airfoil guide vanes 82, at the fan exhaust side 84. Morespecifically, a circumferential row of radially extending airfoil guidevanes 82 are utilized adjacent the fan section 18 to exert somedirectional control of the airflow 78.

Some of the air supplied by the fan 20 can bypass the engine core 44 andbe used for cooling of portions, especially hot portions, of the turbineengine 10, and/or used to cool or power other aspects of the aircraft.In the context of a turbine engine, the hot portions of the engine arenormally downstream of the combustor 30, especially the turbine section32, with the HP turbine 34 being the hottest portion as it is directlydownstream of the combustion section 28. Other sources of cooling fluidcan be, but are not limited to, fluid discharged from the LP compressor24 or the HP compressor 26.

It will be appreciated that the turbine engine 10, and its componentsdescribed herein can be implemented in other turbine engines, includingbut not limited to turbojet, turboprop, turboshaft, and turbofanengines. For example, the turbine engine 10 can be a vanelesscounter-rotating turbine (CRT) engine where both the outer rotor/stator63, and the inner rotor/stator 51 can be rotors.

FIG. 2 is a schematic, partial cross-sectional diagram of an airfoilassembly 102. The airfoil assembly 102 can include a platform 110, adovetail 122, and an airfoil 104 that can be defined by a profile 124.The airfoil 104 can be any airfoil such as a blade or vane in the fansection 18, compressor section 22 or turbine section 32 as desired. Itwill be understood that the airfoil assembly 102 can also include anysuitable component within the turbine engine, including a shroud,hanger, strut, platform, inner band, or outer band, in non-limitingexamples.

For the purposes of explaining this drawing, a coordinate system hasbeen placed wherein the X-axis can be defined by the centerline 12 ofthe turbine engine 10, the Y-axis can be defined by the circumferentialaxis of the turbine engine 10, the Z-axis can be defined by the radialaxis of the turbine engine 10.

The airfoil 104 includes an outer wall 120, which defines a pressureside 130 and a suction side 132. The outer wall 120 can extend aroundthe entirety of the exterior of the airfoil 104 along the pressure side130 and the suction side 132. As such, the outer wall 120 extendsbetween a leading edge 112 and a trailing edge 114 to define achord-wise direction Cd, and also extends in the Z-axis radially betweena root 108 and a tip 106 to define a span-wise direction Sd. A span Scan be defined as the total length of the airfoil in the span-wisedirection Sd from the root 108 to the tip 106.

The airfoil assembly 102 can also include the platform 110 coupled tothe airfoil 104 at the root 108. In one example the airfoil 104 can be ablade of the turbine engine 10. The platform 110 can be mounted to arotating structure to rotate about the X-axis, which rotates the airfoilabout the X-axis. Alternatively, the platform 110 can be mounted to anon-rotating structure, resulting in a non-rotating airfoil. A dovetail122 can depend from the platform 110. In such a case, the platform 110can form at least a portion of the dovetail 122.

Multiple airfoil assemblies 102 can be circumferentially arranged aboutthe X-axis in abutting relationship. The dovetail 122 can be received ina rotating or stationary disk to affect the circumferential arrangement.

The platform 110 shown can be a continuous, unbroken surface.Alternatively, there can be holes, channels, ducts, cracks, troughs, orany other known feature placed throughout the platform 110. Thesevarious exemplarity features of the platform can be used for variousreasons to improve overall engine efficiency. These features can be usedas dust escape, cooling holes, or aerodynamic efficiency boosters.

The dovetail 122 can be configured to mount to at least a portion of theinner rotor/stator 51, or outer rotor/stator 63 of the turbine engine10. The dovetail 122 can comprise a set of inlet passages 116,exemplarily shown as three inlet passages, extending through thedovetail 122 to provide a path for communication for a fluid flow 100 toenter the airfoil 104. Alternatively, there can be any number of inletpassages 116 passing through the dovetail 122 to provide internal fluidcommunication with the airfoil 104. It should be understood that thedovetail 122 is shown in cross-section, such that the inlet passages 116are housed within the body of the dovetail 122.

A plurality of outlets 118 can extend proximate the outer wall 120. Theoutlets 118 are illustrated as being placed in various locations alongthe outer wall 120. The outlets 118 can be placed along the leading edge112, the trailing edge 114, at the root 108 of the airfoil 104, or nearthe tip 106 of the airfoil 104. The outlets 118 can be placed on theouter wall 120 of the pressure side 130 or the suction side 132. Therecan be any number of outlets 118. There can be a plurality of outlets118 where all of the outlets 118 are of the same size and shape. Theoutlets 118 can be of various sizes. For example, the outlets 118 can beas small as a β_(in), or as large as the total length of the airfoil 104in the span-wise direction Sd or the chord-wise direction Cd.

The outlets 118 are illustrated as circular ejection holes. The outlets118 can further include in-line diffusers, diffusing slots, bleed slots,film holes, or channels, in non-limiting examples. While illustrated asbeing circular, the outlets 118 can also have any suitable geometricprofile, including oval, square with rounded corners, orasymmetric/irregular, in non-limiting examples. The outlet 118 can be acontinuous hole or slot leading into the interior of the airfoil incommunication with the fluid flow 100. Alternatively, the outlet 118 caninclude other components such as a porous material, solid material,film, net, and/or any other reasonable material. These other componentscan be placed at or near the outlet 118.

The airfoil 104 can be defined by profile 124 that can create improvedefficiency characteristics. For example, the profile 124 of the airfoil104 can create a preferred contraction ratio (CR), blade turning (BT),solidity, or aspect ratio (AR). The profile 124 can be the same at theroot 108 as it is at the tip 106. Alternatively, the profile of the root108 can be different than that of the tip 106. The entire span S of theairfoil 104 can have the same profile if a cross section were takenalong the rotational X-axis. Alternatively, portions of the airfoil 104can have different profiles 124. For example, if a cross sectional viewwere taken at a first location, for example halfway between the root 108and the tip 106, and a second location, for example at or near the tip106, the profile seen at the first location can be different than theprofile seen at the second location.

FIG. 3 is a cross sectional view of the airfoil 104, and moreparticularly of the profile 124. The profile 124 can be defined withreference to well-known terms for defining airfoils. For example, a meancamber line 126, extending from the leading edge 112 to the trailingedge 114, is a line that is equidistant from the pressure side 130 ofthe outer wall 120 and the suction side 132 of the outer wall 120.

A chord line 128 can be defined as the straight-line distance from theleading edge 112 to the trailing edge 114. As illustrated, for a highlychambered airfoil as shown, the majority of the chord line 128 does notlay within the profile 124 itself, but instead extends through thepressure side 130 area of the airfoil 104. Alternatively, none of, allof, or any fraction between of the chord line 128 may lay within theprofile 124.

The profile 124 includes an inlet angle β_(in) and an outlet angleβ_(out). The inlet angle β_(in) can be defined by the included anglebetween the X-axis and a line parallel to the mean camber line 126 at aleading edge intersection. Similarly, the outlet angle β_(out) can bedefined by the included angle between the X-axis and a line parallel tothe mean camber line 126 at a trailing edge intersection. The leadingedge intersection and the trailing edge intersection can be defined bythe point of intersection of the mean camber line 126 and the leadingedge 112 or trailing edge, respectively. The inlet angle β_(in) can bepositive with respect to the X-axis, while the outlet angle β_(out) canbe negative with respect to the X-axis.

It has been found that the airfoil 104 with the profile 124, which isthe subject of this disclosure, can utilize geometric characteristicssuch as a high contraction ratio (CR), and a low blade turning (BT)along the span S to quantify the profile 124.

The CR of the airfoil 104 can be defined as the geometricalrepresentation of the profile 124, that can produce flow accelerationsfrom the leading edge to the trailing edge. CR can be defined throughthe use of the following equation:

${CR} = {1 - ( \frac{\cos( \beta_{out} )}{\cos( \beta_{in} )} )}$

The CR value of the airfoil 104 can be relatively high when compared toother airfoils, as such, the CR can be defined as a high accelerationcharacteristic of the airfoil 104. The CR value of the airfoil 104 canvary across the span S of the airfoil 104. For example, the CR value canbe greater than 0.55 along at least 80% of the span S, and greater than0.45 between 80% and 100% of the span S. The maximum CR value can occurat midspan defined as a location on the airfoil 104 equidistant from theroot 108 to the tip 106, while the minimum can occur at the root 108 ofthe airfoil 104.

The blade turning, BT of the airfoil 104, can be generally defined as arepresentation of the fluid flow hitting the leading edge 112 of theprofile 124 at an inlet angle β_(in), and then following the curvatureof the outer wall 120 of the profile 124 to the trailing edge 114 whereit exits at an outlet angle β_(out). The total amount of “turning” thefluid flow goes through from the leading edge 112 to the trailing edge114 can be defined as the total BT. BT can be defined through the use ofthe following equation:

${BT} = {{❘\beta_{out}❘} + \frac{\beta_{in}*\beta_{out}}{❘\beta_{out}❘}}$

In the case of the airfoil 104 defined by the profile 124, the BT alongat least a portion of the span S of the airfoil 104 is relatively lowwhen compared to known airfoils, which translates into little curvatureor camber. The BT value can vary along the span S of the airfoil 104.For example, the BT value can be less than 110 degrees over 100% of thespan S, less than 100 degrees between 30% and 50% of the span S, andless than 90 degrees over at least 30% of the span S.

The maximum BT value can occur at a radially inward location of theairfoil 104 and a minimum can occur at a radially outward location ofthe airfoil 104. As used herein, the radially inward location can be aportion of the airfoil 104 that is nearest the centerline 12 of theturbine engine 10, while the radial outward location can be a portion ofthe airfoil 104 farthest the centerline 12. For example, the airfoil 104can be included on the inner rotor 51, in this case, the radially inwardlocation can be the root 108 of the airfoil 104, and the radiallyoutward location can be the tip 106 of the airfoil 104. Conversely, theairfoil 104 can be included on the outer rotor 63 such that the radiallyinward location can be the tip 106 of the airfoil 104, and the radiallyoutward location can be the root 108 of the airfoil 104. Alternatively,one or more portions of the airfoil 104 can have the same BT value.

The airfoil 104 defined by the profile 124 can have a high CR that canfall within the above-mentioned ranges, and a low BT along at least aportion of the span that can fall within the above-mentioned ranges. Thecombination of at least the two can retard boundary layer growth betweenthe air flow and the outer wall 120 of the airfoil 104 on both thepressure side 130 and the suction side 132 from the leading edge 112 tothe trailing edge 114. The retardation of the boundary layer can bebeneficial in ensuring that the fluid flowing around the outer wall 120of the airfoil 104 does not separate too much and create eddies orturbulence along or near the outer wall 120 of the airfoil 104. When alarge amount of turbulence is experienced, pressure losses can begenerated and the overall turbine engine 10 efficiency can bedrastically impacted.

FIG. 4 is a top-down circumferential view of a portion of an annularcomponent 134 of the turbine engine 10 illustrating two airfoils 104spaced about the circumferential Y-axis of the engine.

The effective axial length of each of the airfoils 104 can be defined byan axial chord length C_(ax) extending along the rotational X-axis. TheC_(ax) is the distance between a first radial line from the X-axis thatintersects the leading edge and a second radial line from the X-axisthat intersects the trailing edge. C_(ax) can be thought of as theX-axis component of the chord line. The axial chord length C_(ax) canvary depending on the location of the airfoil 104 in the turbine engine10. For example, the axial chord length C_(ax) can be larger on theannular component 134 at a first location of the turbine engine 10 thananother annular component 134 at second location of the enginecomponent. The axial chord length C_(ax) can be the same for each of theairfoils 104 on a corresponding annular component 134. Alternatively,the axial chord length C_(ax) can vary from each airfoil 104 on acorresponding annular component 134.

The leading edges 112 of the airfoils 104 can be spaced a distance apartdefined by a pitch P extending along the circumferential Y-axis. Pitch Pbetween adjacent airfoils 104 can be constant throughout the annularcomponent 134. In some instance, pitch P is not be constant betweenadjacent airfoils 104. For example, there can be 4 airfoils where thereis a first pitch P1 of a first value between a first and a secondairfoil, a second pitch P2 of a second value between the second airfoiland a third airfoil, a third pitch P3 of a third value between the thirdairfoil and a fourth airfoil, and a fourth pitch P4 of a fourth valuebetween the fourth airfoil and the first airfoil. Each of the first,second, third, fourth, and fifth values can be different. Alternatively,the first and the second values can be the same, while the third, andfourth values are different. All of the first, second, third, and fourthvalues can be the same. It will be appreciated that multiplecombinations can exist.

The airfoil 104 as described herein can greatly reduce the solidity ofthe turbine engine 10. Solidity can be defined as the axial chord lengthC_(ax) over the pitch P. It can be related to the number of profiles 124and hence it can be directly related to the portion of the wetted areawith respect to the fluid flow in the turbine engine 10. The soliditycan be illustrated through the use of the following equation:

${Solidity} = \frac{C_{ax}}{P}$

In this case, the solidity can be relatively low when compared to knownairfoils. The solidity of the turbine engine 10 can be between 0.6 to1.2. Specifically, the solidity of the turbine engine 10 can be 0.7 to0.9.

Reducing the solidity, or wetted area, can provide the advantage ofreducing pressure losses and hence increase overall turbine performance.Additionally, as there can be a lower number of airfoils 104 with alower solidity, the overall airfoil count can be strongly reduced whichcan results in a large weight and cost decrease for the turbine engine10.

FIG. 5 is a side view the airfoil 104 in communication with an annularcomponent 134. The annular component 134 can be any suitable componentadapted to rotate about the rotational X-axis. For example, the annularcomponent 134 can be any portion of the inner rotor 51, or outer rotor63. Alternatively, the annular component 134 can be a bearing, a screw,a cylinder, a gear, or any other known object that can rotate about arotational X-axis.

The airfoil 104 as described prior, can have a greatly reduced AspectRatio (AR) due to its span S and axial chord length C_(ax). The AR canbe defined as the ratio between the profile height, or the span S, ofthe airfoil 104 and the axial chord length C_(ax). A high AR can resultin a relatively tall and narrow airfoil with a substantially rectangularcross section, while a low AR can result in a stubby, or short, airfoilwith a more squared cross section. The AR can be defined through the useof the following equation

${AR} = \frac{S}{C_{ax}}$

The AR of the airfoil 104 can be relatively low when compared to knownairfoils. The AR of the turbine engine 10 can be 2 to 6. Specifically,the AR of the turbine engine 10 can be 3 to 5.

The airfoil 104 can be defined by the profile 124. The profile 124 cancause the airfoil 104 to have a high CR, and low BT at the radiallyinward location can enable a turbine engine 10 with airfoils 104 havinglow AR designs when compared to other airfoils having a lower CR and ahigher BT. The airfoils 104 can have a lower AR which can allow for alower overall number of airfoils 104 needed for the turbine engine 10for a given solidity. This can be beneficial for performance as it canincrease the Reynolds number of the airfoil 104, and reduce the cost ofthe turbine engine 10 as there can be a smaller amount of material usedfor the manufacturing of the turbine engine 10, and the airfoils 104.

As used herein, the terms “high” and “low” are used as a comparison topast airfoil designs. For example, past airfoil designs can have CRvalues ranging from 0.4 to 0.5 at midspan, and 0.2 to 0.3 at a minimumlocation. In comparison, the airfoil 104 defined by the profile 124 canhave a relatively “high” CR value of greater than 0.55 along at least80% of the span S and greater than 0.45 between 80% and 100% of the spanS.

The airfoil 104 can be adapted for use in CRT engines as the relativevelocity of the turbine engine 10 is relatively high when compared to atraditional turbine engine with a single rotor and a single stator perstage. For example, a traditional turbine for a large commercial enginecan rotate with speeds up to 8000 RPM while a CRT engine can rotate withrelative velocities upwards of 12000 RPM. Known airfoils can malfunctionat the higher RPM that can be experienced in CRT engines. The currentairfoil 104 can have a high CR and a low BT at the radially inwardlocation which results in an airfoil better suited for a larger range ofrelative velocities.

The airfoil 104 can be better suited to withstand higher centrifugalforces. During operation of the turbine engine 10, the airfoils 104 canexperience higher centrifugal forces depending on the location of theairfoil in the compressor or the turbine. The airfoil 104 can have ahigher CR, a lower BT at the radially inward location, and a lower ARthat can allow them to withstand higher centrifugal forces than otherknown airfoils. In return, the airfoils 104 can withstand the higherrelative velocities as outlined above.

The airfoil 104 that can have a high CR and a low BT at the radiallyinward location that can allow for a low AR and a low solidity. A low ARand a low solidity can not only decrease losses and hence increaseoverall engine efficiency; however, they can also greatly reduce thenumber of airfoils and the overall weight of the airfoils which greatlyreduces the overall cost of the turbine engine 10.

This written description uses examples to describe aspects of thedisclosure described herein, including the best mode, and also to enableany person skilled in the art to practice aspects of the disclosure,including making and using any devices or systems and performing anyincorporated methods. The patentable scope of aspects of the disclosureis defined by the claims, and can include other examples that occur tothose skilled in the art. Such other examples are intended to be withinthe scope of the claims if they have structural elements that do notdiffer from the literal language of the claims, or if they includeequivalent structural elements with insubstantial differences from theliteral languages of the claims.

Further aspects of the invention are provided by the subject matter ofthe following clauses:

1. A turbine engine comprising at least one blade carried by a rotor androtating about a rotational axis, the blade comprising, an outer walldefining a pressure side and a suction side extending in a chord-wisedirection between a leading edge to a trailing edge and extending in aspan-wise direction between a root and a tip, a mean camber lineextending between a leading edge and a trailing edge and intersectingthe leading edge to define a leading edge intersection, and intersectingthe trailing edge to define a trailing edge intersection, an inletangle, β_(in), defined by an included angle between a line parallel tothe mean camber line at the leading edge intersection and the rotationalaxis, an outlet angle, β_(out), defined by an included angle between aline parallel to the mean camber line at the trailing edge intersectionand the rotational axis, wherein the blade has a contraction ratio (CR)of greater than 0.45 along at least 80% of a span of the at least oneblade, where the CR is determined by the formula:

${CR} = {1 - ( \frac{\cos( \beta_{out} )}{\cos( \beta_{in} )} )}$

wherein the blade has a blade turning (BT) of less than 100 degreesalong at least 30% of the span, where the blade turning is determined bythe formula:

${BT} = {{❘\beta_{out}❘} + \frac{\beta_{in}*\beta_{out}}{❘\beta_{out}❘}}$

2. The turbine engine of any preceding clause wherein β_(in) is constantalong the span from the root to the tip.

3. The turbine engine of any preceding clause wherein β_(out) isconstant along the span from the root to the tip.

4. The turbine engine of any preceding clause wherein CR is greater than0.55 along at least 80% of the span.

5. The turbine engine of any preceding clause wherein a maximum CR valueoccurs at a midspan of the blade between the root and the tip.

6. The turbine engine of any preceding clause wherein BT is less than 90degrees along at least 30% of the span.

7. The turbine engine of any preceding clause wherein BT is less than110 degrees along the entire span.

8. The turbine engine of any preceding clause wherein an aspect ratio ofthe at least one blade is less than 6.

9. The turbine engine of any preceding clause wherein the aspect ratiois less than 5 and at least 3.

10. The turbine engine of any preceding clause wherein the at least oneblade comprises multiple blades that are circumferentially spaced aboutthe rotor.

11. The turbine engine of any preceding clause wherein a solidity isless than 0.9.

12. The turbine engine of any preceding clause wherein an aspect ratioof the multiple blades is less than 5.0.

13. The turbine engine of any preceding clause wherein the CR is greaterthan 0.45 along at least 80% of the span, the BT is less than 100degrees along at least 30% of the span, an aspect ratio is less than 5,and a solidity that is less than 0.9.

14. The turbine engine of any preceding clause wherein the rotor has twocounter-rotating portions, with the at least one blade being carried byat least one of the two counter-rotating portions.

15. The turbine engine of any preceding clause wherein the at least oneblade comprises a first blade on one of the counter-rotating portionsand a second blade on the other of the counter-rotating portions.

16. The turbine engine of any preceding clause wherein the outer wallbounds an interior having at least one cooling air passage, and at leastone cooling hole extends from cooling air passage to an outer surface ofthe outer wall.

17. An airfoil configured to rotate about a rotational axis andcomprising an outer wall defining a pressure side and a suction sideextending in a chord-wise direction between a leading edge to a trailingedge and extending in a span-wise direction between a root and a tip, amean camber line extending between a leading edge and a trailing edgeand intersecting the leading edge to define a leading edge intersection,and intersecting the trailing edge to define a trailing edgeintersection, an inlet angle, β_(in), defined by an included anglebetween a line parallel to the mean camber line at the leading edgeintersection and the rotational axis, an outlet angle, β_(out), definedby an included angle between a line parallel to the mean camber line atthe trailing edge intersection and the rotational axis, wherein theairfoil has a contraction ratio (CR) of greater than 0.45 along at least80% of a span of the at least one blade, where the CR is determined bythe formula:

${CR} = {1 - ( \frac{\cos( \beta_{out} )}{\cos( \beta_{in} )} )}$

wherein the airfoil has a blade turning (BT) of less than 100 degreesalong at least 30% of the span, where the blade turning is determined bythe formula:

${BT} = {{❘\beta_{out}❘} + \frac{\beta_{in}*\beta_{out}}{❘\beta_{out}❘}}$

18. The airfoil of any preceding clause wherein β_(in) is constant alongthe span the root to the tip.

19. The airfoil of any preceding clause wherein β_(out) is constantalong the span from the root to the tip.

20. The airfoil of any preceding clause wherein CR is greater than 0.55along at least 80% of the span.

21. The airfoil of any preceding clause wherein a maximum CR valueoccurs at a midspan of the airfoil between the root and the tip.

22. The airfoil of any preceding clause wherein BT is less than 90degrees along at least 30% of the span.

23. The airfoil of any preceding clause wherein BT is less than 110degrees along the entire span.

24. The airfoil of any preceding clause wherein an aspect ratio is lessthan 5 and at least 3.

25. The airfoil of any preceding clause wherein there are multiplecircumferentially spaced airfoils about a rotor.

26. The airfoils of any preceding clause wherein a solidity is less than0.9.

27. The airfoils of any preceding clause wherein the CR is greater than0.45 along at least 80% of the span, the BT is less than 100 degreesalong at least 30% of the span, an aspect ratio is less than 5, and asolidity that is less than 0.9.

What is claimed is:
 1. An airfoil assembly comprising: at least twoairfoils carried by a rotor and configured to rotate about a rotationalaxis, the at least two airfoils being circumferentially spaced from eachother, with respect to the rotational axis, to define a pitch (P)therebetween, each airfoil of the at least two airfoils comprising: anouter wall defining a pressure side and a suction side extending in achord-wise direction between a leading edge to a trailing edge andextending in a span-wise direction between a root and a tip; an axialchord length (C_(ax)) extending between the leading edge and thetrailing edge in an axial direction with respect to the rotational axis;and an airfoil solidity (Sl) of the at least two airfoils is determinedby: $\frac{C_{ax}}{P};$ wherein the airfoil solidity (Sl) is greaterthan or equal to 0.6 and less than or equal to 1.2 (0.6≤Sl≤1.2).
 2. Theairfoil assembly of claim 1, wherein the airfoil solidity (Sl) isgreater than or equal to 0.7 and less than or equal to 0.9 (0.7≤Sl≤0.9).3. The airfoil assembly of claim 1, wherein each airfoil furthercomprises: a mean camber line extending between the leading edge and thetrailing edge and intersecting the leading edge to define a leading edgeintersection, and intersecting the trailing edge to define a trailingedge intersection; an inlet angle, β_(in), in degrees defined by anincluded angle between a line parallel to the mean camber line at theleading edge intersection and the rotational axis; and an outlet angle,β_(out), in degrees defined by an included angle between a line parallelto the mean camber line at the trailing edge intersection and therotational axis.
 4. The airfoil assembly of claim 3, wherein at leastone of the at least two airfoils has a contraction ratio (CR) of greaterthan 0.45 along at least 80% of a span of the airfoil, where thecontraction ratio (CR) is determined by:${CR} = {1 - ( \frac{\cos( \beta_{out} )}{\cos( \beta_{in} )} )}$5. The airfoil assembly of claim 3, wherein the airfoil has a bladeturning (BT) of less than 100 degrees along at least 30% of the span,wherein the blade turning (BT) is determined by:${BT} = {{❘\beta_{out}❘} + \frac{\beta_{in}*\beta_{out}}{❘\beta_{out}❘}}$6. The airfoil assembly of claim 3, wherein β_(in) is constant along thespan from the root to the tip, or β_(out) is constant along the spanfrom the root to the tip.
 7. The airfoil assembly of claim 1, whereinthe airfoil assembly is configured to be used within a turbine engine.8. The airfoil assembly of claim 7, wherein the turbine engine is acounter-rotating turbine engine and the rotor is one of either an innerrotor or an outer rotor.
 9. An airfoil operably coupled to a rotor andconfigured to rotate about a rotational axis and comprising: an outerwall defining a pressure side and a suction side extending in achord-wise direction between a leading edge to a trailing edge andextending in a span-wise direction between a root and a tip to define aspan (S); an axial chord length (C_(ax)) extending between the leadingedge and the trailing edge in an axial direction with respect to therotational axis; and an aspect ratio (AR) of the airfoil is determinedby: $\frac{C_{ax}}{S};$ wherein the aspect ratio (AR) is greater than orequal to 2 and less than or equal to 6 (2≤AR≤6).
 10. The airfoil ofclaim 9, wherein the aspect ratio (AR) is greater than or equal to 3 andless than or equal to 5 (3≤AR≤5).
 11. The airfoil of claim 9, whereineach airfoil further comprises: a mean camber line extending between theleading edge and the trailing edge and intersecting the leading edge todefine a leading edge intersection, and intersecting the trailing edgeto define a trailing edge intersection; an inlet angle, β_(in), indegrees defined by an included angle between a line parallel to the meancamber line at the leading edge intersection and the rotational axis;and an outlet angle, β_(out), in degrees defined by an included anglebetween a line parallel to the mean camber line at the trailing edgeintersection and the rotational axis.
 12. The airfoil of claim 11,further comprising a contraction ratio (CR) of greater than 0.45 alongat least 80% of a span of the airfoil, where the contraction ratio (CR)is determined by:${CR} = {1 - ( \frac{\cos( \beta_{out} )}{\cos( \beta_{in} )} )}$13. The airfoil of claim 11, wherein the airfoil has a blade turning(BT) of less than 100 degrees along at least 30% of the span, whereinthe blade turning (BT) is determined by:${BT} = {{❘\beta_{out}❘} + \frac{\beta_{in}*\beta_{out}}{❘\beta_{out}❘}}$14. The airfoil of claim 11, wherein β_(in) is constant along the spanfrom the root to the tip, or β_(out) is constant along the span from theroot to the tip.
 15. The airfoil of claim 9, wherein the airfoil isconfigured to be used within a turbine engine.
 16. The airfoil of claim15, wherein the turbine engine is a counter-rotating turbine engine andthe rotor is one of either an inner rotor or an outer rotor.
 17. Anairfoil operably coupled to a rotor and assembly: at least one airfoilconfigured to rotate about a rotational axis, the at least one airfoilcomprising: an outer wall defining a pressure side and a suction sideextending in a chord-wise direction between a leading edge to a trailingedge and extending in a span-wise direction between a root and a tip; amean camber line extending between the leading edge and the trailingedge and intersecting the leading edge to define a leading edgeintersection, and intersecting the trailing edge to define a trailingedge intersection; an inlet angle, β_(in), in degrees defined by anincluded angle between a line parallel to the mean camber line at theleading edge intersection and the rotational axis; an outlet angle,β_(out), in degrees defined by an included angle between a line parallelto the mean camber line at the trailing edge intersection and therotational axis; and a contraction ratio (CR) of greater than 0.55 alongat least 80% of a span of the at least one airfoil, wherein thecontraction ratio (CR) is determined by:${{CR} = {1 - ( \frac{\cos( \beta_{out} )}{\cos( \beta_{in} )} )}};$wherein the contraction ratio (CR) is greater than 0.55 along at least80% of the span.
 18. The airfoil assembly of claim 17, wherein the atleast one airfoil is included within at least two airfoils with eachairfoil of the at least two airfoils comprising: an axial chord length(C_(ax)) extending between the leading edge and the trailing edge in anaxial direction with respect to the rotational axis; a blade turning(BT) of less than 100 degrees along at least 30% of the span, whereinthe blade turning (BT) is determined by:${BT} = {{❘\beta_{out}❘} + \frac{\beta_{in}*\beta_{out}}{❘\beta_{out}❘}}$an airfoil solidity (Sl) of the at least two airfoils is determined by:$\frac{C_{ax}}{P},$  with the airfoil solidity (Sl) being greater thanor equal to 0.6 and less than or equal to 1.2 (0.6≤Sl≤1.2); and anaspect ratio (AR) of the at least one airfoil is determined by:$\frac{C_{ax}}{S},$  with the aspect ratio (AR) being greater than orequal to 2 and less than or equal to 6 (2≤AR≤6).
 19. The airfoilassembly of claim 17, wherein the airfoil assembly is configured to beused within a turbine engine.
 20. The airfoil assembly of claim 19,wherein the turbine engine is a counter-rotating turbine engine and therotor is one of either an inner rotor or an outer rotor.